Variations in the density of a compressible medium (e.g., a gas such as air), known as density inhomogeneities, are important phenomena in aerodynamics. Shocks are one type of density inhomogeneity that can occur when an object moves through the medium faster than the speed at which waves (e.g., sound or pressure waves in air) propagate in the medium. Shocks are regions in the medium characterized by sudden changes in pressure, temperature, and density.
Aerodynamic shocks are usually characterized by parameters such as the shock strength and the shock thickness. The shock strength is conventionally defined as a pressure ratio across the shock or, because it may also be defined as a density ratio, a ratio of refractive indices of the air before and after the shock. The shock thickness is the distance in the direction normal to the shock over which the change in air density occurs. The shock thickness depends upon the mean free path that gas molecules travel before colliding.
Shocks are particularly relevant in the study, design, testing, and performance of vehicles and other objects travelling at supersonic speeds (i.e., having a Mach number between one and five, meaning that the vehicle or object's speed is between one and five times the speed of sound) and hypersonic speeds (i.e., having a Mach number greater than five). In supersonic and hypersonic vehicles, fore-body and engine inlet performance are critical to the overall integrated performance of these vehicles. For instance, scramjet inlets and vehicle fore-bodies are designed for a certain Mach number (shock-on-lip), and their performance deteriorates in other regimes.
Because of the importance of shocks, numerous studies have been performed on shocks and their effects on performance of supersonic vehicles and their components. Various schemes to analyze high speed flows and evaluate performance of aircraft and spacecraft components have been developed. A need for shock position sensors capable of meeting flight qualifying requirements has been recognized and attempts have been made to develop such sensors. Early efforts were concentrated around using pressure taps along the inlet walls. The positions of the shocks were determined by tracking the pressure reading and locating the pressure jump associated with the shock. This basic technique evolved into several wall pressure-based configurations of normal shock position sensing systems. Despite apparent initial success, these wall pressure-based measuring techniques have serious drawbacks. Two important drawbacks are slow response due to pneumatic manifolds used and the effect of the boundary layer on the stability of pressure readings. These issues can seriously restrict applicability of these techniques to normal shock detection and control during supersonic flight.
Moreover, for a commercial aircraft, economic efficiency has to be achieved in order to make supersonic flight economically viable. As a result, an effective control system is required, in addition to avoiding an unstart, to provide the most economical operating regime for the engine (achieved by minimizing the fuel consumption).
Optical flow analyzing methods do not have the same issues of the wall pressure-based measuring techniques and optical flow visualization is widely used in ground-based flow analyzing facilities. Effects of propagation of light through density inhomogeneities have been conventionally detected and visualized by interferometers, Schlieren systems, and shadowgraphs.
Flow visualization techniques such as interferometric, Schlieren, and shadowgraphy typically involve a laser or other source of light and a collimating lens that forms a nearly plane wave. The wave is sent through a transparent section of the test facility normal to the direction of the air flow. After passing through the transparent section of the facility, the plane wave is displayed on a screen or a charge-coupled device (CCD) or diode array. If the air flow is homogeneous and the air density is constant everywhere inside the test section, the display is uniformly illuminated. However, if the flow contains density variations, the illumination of the display is not uniform but rather has dark and bright regions. The contrast of the resultant pattern depends on the strength of the density variations or the density gradient as well as the visualization technique used. Among conventional flow visualization techniques, e.g., interferometry, Schlieren, and shadowgraph, the shadowgraph is often considered one of the most suitable for shock detection. It is because patterns generated by the technique represent the second order derivatives of the density distribution, and the shocks that are being created by very rapid changes in air densities are traditionally observed best by the shadowgraph.
These conventional flow visualization techniques, despite their wide use, have significant drawbacks. First of all, the techniques are based on filling most of the window of the test section and require high power light sources and large optical components. Thus, they cannot be economically or efficiently used in air- or space-borne systems without significant weight and real estate penalties. On the other hand, small and lightweight low power light sources in the conventional configuration do not generate a signal with a sufficient signal-to-noise ratio at the detector to achieve an adequate resolution. Secondly, the fact that the entire test section has to be illuminated masks the second order phenomena associated with the wave propagation through and interaction with inhomogeneities.
As vehicles are developed with speeds increasingly approaching hypersonic regimes, formation of shocks and their interaction gain even more importance and the need to develop in-flight shock sensing and mitigation technology becomes even more acute. However, the space and weight requirements of conventional systems make them untenable for use on such vehicles.